Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices was designed and tested. The thrust stand was specifically tailored to the needs of a 100 to 250 kW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. A detailed description is given for the thrust stand design and operation with a 100 kW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability

Arcjet component conditions through a multistart test

A low power, dc arcjet thruster was tested for starting reliability using hydrogen-nitrogen mixtures simulating the decomposition products of hydrazine. More than 300 starts were accumulated in phases with extended burn-in periods interlaced. A high degree of flow stabilization was built into the arcjet and the power supply incorporated both rapid current regulation and a high voltage, pulsed starting circuit. A nominal current level of 10 A was maintained throughout the test. Photomicrographs of the cathode tip showed a rapid recession to a steady-state operating geometry. A target of 300 starts was selected, as this represents significantly more than anticipated (150 to 240), in missions of 10 yr or less duration. Weighings showed no apparent mass loss. Some anode erosion was observed, particularly at the entrance to the constrictor. This was attributed to the brief period during startup the arc mode attachment point spends in the high pressure region upstream of the nozzle. Based on the results obtained, startup does not appear to be performance or life limiting for the number of starts typical of operational satellite applications

High-power hydrogen arcjet performance

A hydrogen arcjet was operated at power levels ranging from 5 to 30 kW with three different nozzle geometries. Test results using all three nozzle geometries are reported and include variations of specific impulse with flow rate, and thrust with power. Geometric variables investigated included constrictor diameter, length, and diverging exit angle. The nozzle with a constrictor diameter of 1.78 mm and divergence angle of 20 degrees was found to give the highest performance. A specific impulse of 1460 s was attained with this nozzle at a thrust efficiency of 29.8 percent. The best efficiency measured was 34.4 percent at a specific impulse of 1045 s. Post test examination of the cathode showed erosion after 28 hours of operation to be small, and limited to the conical tip where steady state arc attachment occurred. Each nozzle was tested to destruction

Arcjet starting reliability: A multistart test on hydrogen/nitrogen mixtures

An arcjet starting reliability test was performed to investigate one feasibility issue in the use of arcjets onboard a satellite for north-south stationkeeping. A 1 kW arcjet was run on hydrogen/nitrogen gas mixtures simulating decomposed hydrazine. A pulse width modulated power supply with an integral high voltage starting pulser was used for arc ignition and steady-state operation. The test was performed in four phases in order to determine if starting characteristics changed as a result of long term thruster operation. More than 300 successful starts were accumulated over an operating time of 18 hrs. Overall results indicate that there is a link between starting characteristics and long term thruster operation; however, the large number of starts had no effect on steady-state performance

Ion Thruster Support and Positioning System

A system for supporting and selectively positioning an ion thruster relative to a surface of a spacecraft includes three angularly spaced thruster support assemblies. Each thruster support assembly includes a frame which has a rotary actuator mounted thereon. The rotary actuator is connected to an actuator member which is rotatably connected to a thruster attachment member connected to a body of the thruster. A stabilizer member is rotatably mounted to the frame and to the thruster attachment member. The thruster is selectively movable in the pitch and yaw directions responsive to movement of the actuator members by the actuators on the thruster support assemblies. A failure of any one actuator on a thruster support assembly will generally still enable limited thruster positioning capability in two directions. In a retracted position the thruster attachment members are held in nested relation in saddles supported on the frames of the thruster support assemblies. The thruster is securely held in the retracted position during periods of high loading such as during launch of the spacecraft

Ion Thruster Support and Positioning System

A system for supporting and selectively positioning an ion thruster relative to a surface of a spacecraft includes three angularly spaced thruster support assemblies. Each thruster support assembly includes a frame which has a rotary actuator mounted thereon. The rotary actuator is connected to an actuator member which is rotatably connected to a thruster attachment member connected to a body of the thruster. A stabilizer member is rotatably mounted to the frame and to the thruster attachment member. The thruster is selectively movable in the pitch and yaw directions responsive to movement of the actuator members by the actuators on the thruster support assemblies. A failure of any one actuator on a thruster support assembly will generally still enable limited thruster positioning capability in two directions. In a retracted position the thruster attachment members are held in nested relation in saddles supported on the frames of the thruster support assemblies. The thruster is securely held in the retracted position during periods of high loading such as during launch of the spacecraft

Recent testing of 30 kW hydrogen arcjet thrusters

NASA is conducting efforts to evaluate high-power hydrogen arcjets for orbit transfer propulsion applications. As part of this program, an attempt was made to reexamine both radiatively- and regeneratively-cooled, 30 kW thrusters first demonstrated by the Giannini Scientific Corp. in 1963. The arcjets were configured to force arc attachment upstream of the throat in a subsonic chamber region. While thruster currents were steady, the voltage traces exhibited sawtooth waveforms at frequencies on the order of 20 kHz. Voltage variations per cycle were typically between 100 and 310 volts, indicating major changes in the position of the arc attachment with time. When operated at their respective design points, the performance of both thrusters fell below the values listed in the 1960's development reports. The reason for the discrepancies is not currently understood and further investigations are in progress. However, the recently measured efficiencies were high compared to those obtained with constricted-arc designs at similar conditions, and further arcjet performance optimizations may be possible

Arcjet cathode phenomena

Cathode tips made from a number of different materials were tested in a modular arcjet thruster in order to examine cathode phenomena. Periodic disassembly and examination, along with the data collected during testing, indicated that all of the tungsten-based materials behaved similarly despite the fact that in one of these samples the percentage of thorium oxide was doubled and another was 25 percent rhenium. The mass loss rate from a 2 percent thoriated rhenium cathode was found to be an order of magnitude greater than that observed using 2 percent thoriated tungsten. Detailed analysis of one of these cathode tips showed that the molten crater contained pure tungsten to a depth of about 150 microns. Problems with thermal stress cracking were encountered in the testing of a hafnium carbide tip. Post test analysis showed that the active area of the tip had chemically reacted with the propellant. A 100 hour continuous test was run at about 1 kW. Post test analysis revealed no dendrite formation, such as observed in a 30 kW arcjet lifetest, near the cathode crater. The cathodes from both this test and a previously run 1000 hour cycled test displayed nearly identical arc craters. Data and calculations indicate that the mass losses observed in testing can be explained by evaporation

Test Facility and Preliminary Performance of a 100 kW Class MPD Thruster

A 260 kW magnetoplasmadynamic (MPD) thruster test facility was assembled and used to characterize thrusters at power levels up to 130 kW using argon and helium propellants. Sensitivities of discharge characteristics to arc current, mass flow rate, and applied magnetic field were investigated. A thermal efficiency correlation developed by others for low power MPD thrusters defined parametric guidelines to minimize electrode losses in MPD thrusters. Argon and helium results suggest that a parameter defined as the product of arc voltage and the square root of the mass flow rate must exceed 0.7 V/kg(sup 1/2)/sec(sup 1/2) in order to obtain thermal efficiencies in excess of 60 percent

Hydrogen arcjet technology

During the 1960's, a substantial research effort was centered on the development of arcjets for space propulsion applications. The majority of the work was at the 30 kW power level with some work at 1-2 kW. At the end of the research effort, the hydrogen arcjet had demonstrated over 700 hours of life in a continuous endurance test at 30 kW, at a specific impulse over 1000 s, and at an efficiency of 0.41. Another high power design demonstrated 500 h life with an efficiency of over 0.50 at the same specific impulse and power levels. At lower power levels, a life of 150 hours was demonstrated at 2 kW with an efficiency of 0.31 and a specific impulse of 935 s. Lack of a space power source hindered arcjet acceptance and research ceased. Over three decades after the first research began, renewed interest exists for hydrogen arcjets. The new approach includes concurrent development of the power processing technology with the arcjet thruster. Performance data were recently obtained over a power range of 0.3-30 kW. The 2 kW performance has been repeated; however, the present high power performance is lower than that obtained in the 1960's at 30 kW, and lifetimes of present thrusters have not yet been demonstrated. Laboratory power processing units have been developed and operated with hydrogen arcjets for the 0.1 kW to 5 kW power range. A 10 kW power processing unit is under development and has been operated at design power into a resistive load